The orbits and the orientations of satellites and of other exo-atmospheric vehicles, referred to as “EV” hereinbelow, require correction to compensate for the effect of drag and drift, to correct trajectory such as for insertion into orbit and for change of orbit, and for fulfilling further operational requirements. Some of these corrections are carried out by small rocket thrusters, named “thrusters” hereinbelow, using hydrazine or other liquids as propellants, which thrusters are operated whenever a functional need arises. In the following description, hydrazine propellant is discussed, although additional liquid propellants are compatible for use within the scope of the present invention.
Many EVs are equipped with a fairly large number of such thrusters, e.g. twelve. Those thrusters of an EV, which operate in steady mode firing, may be replaced by the thruster with electro-thermal thrust augmentation, for the sake of more efficient use of the propellant. When compared with existing electro-thermal augmented thrusters, the thruster of the present invention presents benefits such as reduced weight, lower production costs and expedient assembly. The saved weight is then available for additional payload and/or for supplementary propellant.
Evidently, there are severe weight and volume limitations for an EV. The amount of propellant carried on board is usually the decisive factor limiting the useful operational life of an EV. Furthermore, the propellant supply of the EV cannot usually be replenished after launch, and hence, once exhausted, the EV can no longer be controlled or maneuvered.
The ratio between the total impulse, divided by the propellant weight used to generate it, is defined as the Specific Impulse, and is designated as Isp. Consequently, the higher the specific impulse, the lower the propellant mass required to provide a given impulse. The Isp depends primarily on the properties of the propellant, but improved thruster designs may effectively augment the obtained Isp. Specific Impulse, or Isp, is measured in N-sec/kg, and for hydrazine, a typical value obtained for a commonly known, unaugmented thruster design is 2220 N-sec/kg.
Hydrazine and other monopropellants differ from other liquid propellants by the use of a catalyst for causing exothermic reaction of the propellant into gaseous products, in contrast with conventional bi-propellant reaction between a fuel and an oxidizer.
The basic structure and operation mode of a hydrazine thruster and of other monopropellant thrusters is described below. A propellant control valve controls duration of propellant supply, while the mass flow rate of the propellant is governed mostly by the inlet pressure of the propellant injected into the thruster's chamber. The injected propellant contacts a preferably pre-heated catalyst and is decomposed exothermally into hot gaseous products, which are then ejected into space through a nozzle. Specifically for hydrazine thrusters, ammonia, which is part of the decomposed products, is further partially dissociated through an endothermic process as it passes through the remainder of the catalyst.
In current design practice, the catalyst is heated by an electric resistive heater assembled as an external unit, referred to below as “catalyst heater”, mounted outside, or on an outside portion of a conventional thruster. The catalyst heater raises the temperature to a range of preferably 100° C. to 200° C., and requires electric power in the order of magnitude of a few Watts.
Known thruster designs take advantage of the fact that most EVs include a solar electric power supply system, whose power output is at times well in excess of the EV's various requirements. The excess power output may thus be applied for the heating of the gaseous products through the use of filaments forming another electric resistor heater, or “heater” hereinbelow, before their ejection out of the thruster and into space, as taught, for example by U.S. Pat. No. 4,305,247 by Ellion et al., and by U.S. Pat. No. 4,569,198 No. by Cann. This additional thermal energy is added to the gaseous products—minus the heat losses—and raises the temperature of the ejected gases, thereby contributing to a higher Isp value. Typically, such a heater provides a power of about 600 to 1000 Watts per Newton of thrust, and may increase the Isp of a thruster by several tens of percents.
U.S. Pat. No. 4,305,247 to Ellion et al. is for augmenting specific impulse by heating gasses, as recited in column 1, lines 6-9: “This invention is directed to a hydrazine thruster which has augumented specific impulse by heating of the gases which are the product of hydrazine dissociation”, which is achieved as recited in column 3, lines 32-34, by use of: “an electrically resistive tubular ceramic resistance heater positioned for heating gas in said heater chamber”.
However, it is further stated, from column 3, line 35, to column 4, line 4 that: “an exterior protector tube positioned around the exterior of said tubular ceramic resistance heater and an interior protector tube positioned within said tubular ceramic resistance heater, said protector tubes being heated by said tubular ceramic resistance heater to prevent hot gas from the decomposition chamber from flowing directly against said tubular ceramic resistance heater, said protector tubes being positioned so that hot gas flows both exteriorly and interiorly thereof”, which clearly describe indirect heating of the gas since an exterior protection tube and an interior protection tube prevent hot gas from flowing in direct contact with the heater.
The present invention utilizes direct heating in contrast with Ellion et al. who use indirect heating, which suffers from heat losses, and is evidently less efficient than direct heating.
U.S. Pat. No. 4,322,946 by Murch et al. claim a heater, or a superheater for heating gasses, as by its title: “Thermal thruster with superheater”. Murch et al. recite in col 3, lines 11-15: “The superheater continuously and steadily imparts energy to gases emitted from the heater chamber to provide a sensibly constant power output with no discontinuity or irregularities. The superheater increases the temperature of the chamber gases”.
More details about the superheater, or superheater section are provided in relation to FIG. 1, as recited from column 3, line 63 to column 4, line 4: “In FIG. 1, one embodiment of a superheater section 17 is provided for the heater chamber 11 and comprises an elongate cylindrical chamber 18 axially aligned with and abutting the heater chamber. The cylindrical chamber 18 terminates in an annulus 19. A hollow, electrically heated, coiled rhenium tube 20 is disposed centrally along chamber 18 and receives decomposed gases from the heater chamber 11 at its outlet 11a; the tube 20 terminates in a nozzle 21 which fits into the chamber annulus 19.”
Murch et al. use vertical flow of gas through a coiled rhenium tube, but not axial flow and ceramic heaters as with the present invention.
U.S. Pat. No. 4,656,828 to Bingley et al. is for a temperature control system, as by its title: “Augmentation heater temperature control system”. It is also stated in column 1, lines 5-6:” This invention relates to a temperature control system for an augmentation heater for a hydrazine thruster”, and is further recited in column 1, lines 31-36: “In accordance with one embodiment of the present invention the augmentation heater wire element is maintained at the same temperature by a control circuit which connects and disconnects the battery bus to and from the heater wire element in accordance with the resistance sensed across the heater wire element”.
U.S. Pat. No. 5,819,526 to Jackson et al. for a propellant feed system is entitled accordingly as “Low power arcjet propellant feed system”. Jackson et al. recite in column 2, lines 31-42: “In accordance with the teachings of the present invention, a low power arcjet propellant feed system for delivering propellant to a low power arcjet is disclosed. The low power arcjet propellant feed system provides a substantially continuous and stable low flow rate of a gaseous propellant to the low power arcjet. This substantially continuous and stable low flow rate enables precision thrust control of the low power arcjet and stable arcjet operation. Moreover, the substantially continuous and stable low flow rate can be controllably adjusted so that the thrust from the low power arcjet can be dynamically varied over a wide range as required”.
To remove possible doubts, there is also recited in column 2, lines 56-61: “Use of the present invention provides a low power arcjet propellant feed system for delivering propellant to a low power arcjet. As a result, the aforementioned disadvantages associated with utilizing the currently available propellant feed systems have been substantially eliminated.”
The prior art thus does not disclose, teach or suggest the subject matter of the present invention.
Owing to the corrosive effects of the hydrazine products on the commonly used resistive heater filaments, it is currently taught by the above-mentioned patents and by many other sources, that any direct contact between the hydrazine products and the hot resistor filament of a heater should be avoided. Direct contact is avoided by placing some kind of a partition, or wall, between the heater and the hydrazine products, thereby preventing direct heat transfer to the hydrazine's gaseous products. However, indirect heating complicates the thruster's structure, adds to its weight, increases response time, and boosts costs. Moreover, indirect heating calls for a larger temperature difference between the heater and the heated gas than would be otherwise necessary, because of the additional thermal resistance penalty inflicted by the partition.
The higher heater temperature also leads to a shorter heater operational life, to higher heat losses, and therefore to higher power requirement for a given performance. In addition, regarding thermal isolation and heat removal of excess heat from the EV, the problems involved become more elaborate, requiring sophisticated and expensive solutions. Evidently, the heat loss from the heater must be removed out of the EV.
The electro-thermally augmented thruster with internal heaters, as by the present invention, overcomes or significantly reduces the aforementioned problems.